Integrated power unit

ABSTRACT

Duplication of engine and secondary power system components in individual auxiliary and emergency power units is avoided in an integrated power unit including a gas turbine engine 10 having a centrifugal compressor 28 and a turbine 40, a generator 18 coupled to the engine 10, an air inlet 22 to the compressor 28, a first reservoir 82 for combustible fuel, and a first combustor 34, 36 for receiving and combusting compressed air from the compressor 28 and fuel from the reservoir 82 to generate gases of combustion and feeding the gases of combustion to the turbine 40 by the provision of a second reservoir 80 for containing a compressed oxidant for combusting fuel from the first reservoir 82, a second combustor 48 for receiving and combusting fuel from the first reservoir 82 and oxidant from the second reservoir 80 to generate gases of combustion, flow paths 64, 66, 68 for introducing fuel from the reservoir 82 into the gases of combustion from the second combustor 48 to vaporize the fuel and a conduit 50, 52 for directing the combined gases of combustion and vaporized fuel to the turbine 40 to drive the same. The apparatus is useful as an auxiliary power unit when the combustor 34, 36 is in use and as an emergency power unit or starter when the combustor 48 is in use.

This is a division of application Ser. No. 001,818 filed Jan. 8, 1987.

Field of the Invention

This invention relates to a power unit useful in, for example, aircraftwhich combine functions of an auxiliary power unit and an emergencypower unit and thus is an integrated power unit

BACKGROUND OF THE INVENTION

In so-called fly-by-wire aircraft, aircraft control surfaces are notlinked to the controls by mechanical means. Rather, the linking is viaelectrical or hydraulic circuits. Consequently, in the event of anelectrical power or hydraulic failure, the aerodynamic configuration ofthe aircraft cannot be altered under the control of the pilot untilpower is restored. As a result, such aircraft require an emergency powerunit which is capable of responding to a power failure and providing asizable quantity of electrical or hydraulic energy in very short orderso that control of the aircraft can be returned to the pilot.

Fly-by-wire aircraft, like other aircraft of more than basic simplicity,also require an auxiliary power unit for providing electrical andhydraulic energy and bleed air when the main engine or engines of theaircraft are not in use.

Quite typically, both an emergency power unit and an auxiliary powerunit will employ a gas turbine engine coupled to a generator and ahydraulic pump. Thus, where an aircraft employs an emergency power unitand an auxiliary power unit, it will have two turbines, two generatorsand two pumps. This of course requires a certain space on the aircraftand will cause some weight concerns.

While in some aircraft an auxiliary power unit may be easily adapted toserve as an emergency power unit as well, the adaptation is not sosimple on high performance aircraft that may operate at rather highaltitudes. In particular, because a typical auxiliary power unit turbineis an air breathing turbine, at high altitudes the density of the airwill be insufficient to start the turbine and rapidly bring the same upto a speed at which it will operate at that altitude to produceemergency power.

To meet these and other problems, Friedrich, in his U.S. Pat. No.4,092,824, issued June 6, 1978, proposes a turbine for use in aircraftfor starting purposes as well as for driving auxiliary equipment such asa generator and which is capable of operating in a conventional airbreathing mode as well as in an emergency mode that does not require thepresence of air. In particular, Friedrich includes a supply of hydrazineon the aircraft Hydrazine is capable of undergoing an exothermicdecomposition reaction. According to Friedrich, the heat from thisreaction is utilized to vaporize aircraft fuel to provide gas to drivethe turbine in an emergency situation.

While the Friedrich solution does solve a number of the previouslyspecified problems, it also creates a few new ones. In particular, thedecomposition products of hydrazine can accumulate much like soot withinthe turbine, something that will decrease turbine efficiency whenoperated conventionally. Perhaps more significantly, because the basisof the Friedrich system is that of an exothermic decomposition reaction,it necessarily follows that a fuel, such as hydrazine, which is utilizedin the system must be sufficiently unstable as to rapidly undergodecomposition. Of course, the presence of a fuel that is not stable inthe conventional sense on an aircraft presents hazards of its own.

Still another difficulty resides in the fact that hydrazine and properstorage facilities therefor may not be available at all locations. Thusservicing of a system whose hydrazine fuel charge has been partially orwholly consumed becomes a problem. In addition, hydrazine is toxic.Consequently, it is not easily handled.

Finally, to operate in the non air breathing mode, Friedrich requiresthe mechanical decoupling of the engine compressor from the turbine.This not only increases the complexity of the engine, but will increaseits size and weight as well.

The present invention is directed to overcoming these problems.

SUMMARY OF THE INVENTION

It is the principal object of the invention to provide a new andimproved integrated power unit, that is, a power unit which combines thefunctions of an auxiliary power unit and an emergency power unit. Moreparticularly, it is an object of the invention to provide such anintegrated power unit wherein readily available, non-toxic, easilystored stable fuels which, when utilized to drive the unit in anemergency situation, do not leave residue within the unit.

An exemplary embodiment of the invention achieves the foregoing objectin an integrated power unit that is alternatively operable as anauxiliary power unit and as an emergency power unit. The integratedpower unit includes a gas turbine engine having a centrifugal compressorcoupled to a turbine. A generator and/or hydraulic pump is coupled tothe turbine engine to be driven thereby and there is an air inlet to thecompressor. The system includes a first reservoir for a combustible fuel(such as aircraft fuel) and a first combustor for receiving andcombusting compressed air from the compressor and fuel from thereservoir to generate gasses of combustion. Those gasses of combustionare fed to the turbine for normal operation of the same as an auxiliarypower unit. The system further includes a second reservoir forcontaining a compressed oxidant for combusting fuel from the firstreservoir. Preferably, the oxidant is a readily available stablematerial such as air or oxygen. A second combustor is provided forreceiving and combusting fuel from the first reservoir and oxidant fromthe second reservoir to generate gasses of combustion. Means areincluded for introducing fuel from the first reservoir into the gassesof combustion from the second combustor to vaporize such fuel and meansare provided for directing combined gasses of combustion and vaporizedfuel to the turbine to drive the same to thereby provide for emergencyoperation.

Because, in the preferred embodiment, in an emergency operation, theturbine is being driven by a combination of vaporized, uncombusted fueland gasses of combustion from oxidation of the regular fuel by anoxidant such as air, the gasses flowing through the turbine duringoperation as an emergency power unit will have similar characteristicsas those flowing through the unit when used as an auxiliary power unit.Thus, operation as an emergency power unit will have no more deleteriouseffect on subsequent turbine performance than operation of the unit asan auxiliary power unit. Furthermore, the use of an oxidant such as airwith the conventional fuel already on board eliminates the addition ofany chemical compound of less than desirable stability in favor of thoserequired by the aircraft in any event. Consequently, the system does notrequire an increase in chemical hazard. Furthermore, because the fuel isjet fuel and where the oxidant is air, both are readily available toavoid logistics difficulty in servicing.

In a preferred embodiment, the air inlet includes variable inlet guidevanes of a conventional construction which are conventionally intendedto control the flow of air from the inlet to the compressor. Preferably,when the apparatus is in the emergency power unit mode, such vanes areemployed to close the inlet. This effectively removes the vast majorityof the load placed on the system during acceleration of the compressorand minimizes the need to provide for decoupling of the compressor fromthe turbine when the apparatus is utilized in the emergency power unitmode.

In flight operation at high altitude the inlet guide vanes furtherminimize the aerodynamic parasitic drag of the compressor and turbine.Thus the invention is particularly suitable for emergency operation athigh altitudes.

In a highly preferred embodiment, the engine includes a turbine exhaustsection and the second combustor is located within the exhaust sectionto conserve space and to take advantage of already present insulationmaterials. The invention also contemplates the second combustor includea combustion chamber surrounded by a fluid jacket with the fluid jacketbeing in fluid communication and interposed between the first reservoirand the introducing means so that the fuel to be vaporized cools thecombustion chamber.

This feature of the invention enables combustion within the secondcombustor during the emergency power unit mode to occur at maximumtemperature to make more thermal energy available for quickly vaporizingfuel to be fed to the turbine as vapor. This in turn increases the powerdensity through the turbine which increases its response for start up aswell as its power generating capability.

The invention also contemplates a method of operating such a turbine togenerate emergency power which includes the steps of using the variableinlet guide vanes to substantially close the inlet while directing fueland compressed oxidant to the second combustor to combust the same anddirecting fuel to the introducing means.

The invention contemplates that such steps are performed substantiallysimultaneously and that the endurance of the performance of the steps issufficient long as to bring the engine up to a speed whereat it mayoperate on gasses of combustion from the first combustor.

In the highly preferred embodiment, the step of directing fuel andcompressed oxidant to the second combustor is performed so that theoxidant and the fuel are combusted approximately stoichiometrically.

Other objects and advantages will become apparent from the followingspecification taken in connection with the accompanying drawings.

DESCRIPTION OF THE DRAWINGS

FIG. 1 is an elevational view of an integrated power unit made accordingto the invention with parts broken away for clarity;

FIG. 2 is an enlarged sectional view of a combustor employed in theinvention;

FIG. 3 is a schematic view of the invention operating in an auxiliarypower unit mode;

FIG. 4 is a schematic view like FIG. 3 but illustrating operation in theemergency power unit mode.

DESCRIPTION OF THE PREFERRED EMBODIMENT

An exemplary embodiment of an integrated power unit made according tothe invention is illustrated in the drawings and with reference to FIG.1 is seen to include a turbine engine, generally designated 10, having ashaft 12 entering a gearbox, generally designated 14. Mounted on thegearbox 14 are accessories such as a pump 16 and an electrical generator18. The generator 18 includes an input shaft 20 geared to the shaft 12within the gearbox 14. Thus, the turbine 10, when operating, will drivethe generator 18 and the pump 16.

The engine 10 includes a conventional air inlet 22 which in turn isprovided with a plurality of variable inlet guide vanes 24, only one ofwhich is shown. The vanes 24 are operable in a conventional fashion forconventional purposes to control the flow of air from the inlet 22 to aconventional radial discharge compressor 28 mounted on the shaft 12.Vane modulation may also be used for increased surge margin.

A diffuser 30 of conventional construction is disposed radiallyoutwardly of the compressor 28 in a housing 32. Compressed air from thecompressor 28, after passing through the diffuser 30 is fed viaconventional flow passages within a combustor 34 to an annularcombustion chamber 36 whereat it is combusted. Gasses of combustion exitthe combustion chamber 36 via a radial nozzle structure 38 which opensradially inwardly to a turbine 40, also on the shaft 12. Thus, thegasses of combustion resulting from combustion within the chamber 36drive the turbine 40 to cause rotation of the shaft 12 which in turnboth drives the compressor 28 and the various accessories including thegenerator 18 via the gearbox 14; and this mode of operation is that ofan auxiliary power unit. If desired the turbine 40, which is illustratedas a single stage turbine, can be a plural stage turbine, having turbinewheels on several shafts.

The engine 10 also includes a turbine exhaust housing 42 of generallyconventional construction save for modification shortly to beidentified. The turbine exhaust housing 42 has an outlet opening 44which is off of the axis of rotation of the shaft 12.

An opening 46 in the housing 42 oppositely of the turbine 40 provides ameans whereby a second combustor, generally designated 48, may bedisposed within the housing 42 for space conservation purposes.Alternatively, the combustor may be located exteriorly of the housing 42at any desired remote location. Also within the housing 42, a conduit 50extends from the outlet of the combustor 48 to a second annular nozzleassembly 52 about the turbine 40. The nozzle assembly 52 providessupersonic nozzles and as a consequence, it will be appreciated that gasunder pressure from the combustor 48 will be applied to the turbine 40to drive the same.

The combustor 48 is shown in greater detail in FIG. 2 and includes acombustion chamber 56 bounded by an interior cylindrical wall 58terminating in a frustoconical wall 60 extending to an outlet 62 whichis connected to the conduit 50. As can be seen, the combustion chamber56 is jacketed by fluid flow passages 64 which extend to outlet openings66 at the outlet 62. An inlet to the passages is provided at 68 and isadapted to receive fuel from the fuel tanks of the aircraft in which theintegrated power unit is to be installed.

The combustor 48 also includes an opening 70 for receipt of an ignitiondevice as well as an inlet 72 connected to an interior nozzle 74 on theaxis of the combustion chamber 56 and through which fuel may beintroduced into the combustion chamber 56. The inlet 72, like the inlet68, is adapted to be connected to the fuel tanks for the aircraft inwhich the integrated power unit is to be installed.

An oxidant inlet chamber 76 is formed in the combustor 48 oppositely theoutlet 72 and in surrounding relation to the nozzle 74. The inletchamber 76 includes an inlet 78 which is adapted to be connected to asource of compressed oxidant as, for example, a storage bottle forcompressed air shown schematically at 80 in FIGS. 3 and 4.

The combustor 48 is thus modular in form and may be installed as amodule within the opening 48 and secured in place by means of threadedfasteners 80 (FIG. 1) only one of which is shown.

Turning now to FIGS. 3 and 4, fuel tanks for the system and for theaircraft for which it is installed are shown schematically at 82.Operation of the apparatus in auxiliary power unit mode is illustratedschematically in FIG. 3. In particular, with the engine 10 in operation,air is drawn through the inlet 22 and past the variable inlet guidevanes 22 in the direction illustrated by arrows 84 to the compressor 28.The same compresses such air which is then fed to the combustor 34, 36.In addition, some part of the compressed air may be drawn off as bleedair if required as indicated by an arrow 86.

Aircraft fuel from the tank 82 is also introduced into the combustor 34,36 and burned therein in a conventional fashion. Combustion gas isultimately discharged through the annular nozzle 38 against the turbine40 as illustrated by arrows 88 to drive the same. That of course willresult in the generator 18 and the pump 16 being driven by thisconnection to the shaft 12 through the gearbox 20. Thus, both electricaland hydraulic power and bleed air are made available during this mode ofoperation. Control of the operation can be made conventionally includingvarying the inlet geometry by varying the positions of the variableinlet guide vanes 24.

When operated as an emergency power unit or as the starter mode for APUstart initiation, system components are generally as illustrated in FIG.4. In this connection, it will be noted that the variable inlet guidevanes 24 have been rotated to close the inlet 22 from the compressor 28.Thus, the compressor 28 will not be performing work in compressing airupon rotation of the shaft 12.

Stored oxidant, preferably air, from the storage bottle 80 is admittedto the combustor 48 via the inlet 78. At the same time, jet fuel fromthe aircraft tank 82 is entering the combustor 48 via the inlet 72. Thetwo are combusted and in the process generate hot gasses of combustion.These hot gasses of combustion are fed by the conduit 50 by the annularnozzle 52. At the same time, fuel from the aircraft tank 82 is beingintroduced into the combustor 48 adjacent its outlet 62 via the flowpath including the inlet 68 and the outlet openings 66. This fuelintroduction is downstream of the combustion zone within the combustor48 as will be readily appreciated from FIG. 2 and the fuel thusintroduced will not burn. Rather, it will cool the gasses of combustionas well as the walls of the combustion chamber 56. In the process, suchfuel will be vaporized and will flow to the annular nozzle structure 52via the conduit 50. Thus, the combination of the hot gasses ofcombustion provided as the result of combustion within the combustor 48as well as vaporized fuel now impinge upon the turbine 40 to drive thesame to ultimately drive the generator 18 and pump 16. Preferably,oxidant and fuel fed into the combustor 48 are in the correct ratios sothat the combustion reaction occurs stoichiometrically. This in turnmaximizes the temperature of the gasses thus generated and the heatavailable from such combustion to thereby maximize the amount of fuelthat can be vaporized as it emanates from the outlet openings 66. Thisin turn means that the mass flow rate of gas from the combustor 48through the turbine 40 is maximized. Thus, upon start up in either theAPU start mode or the emergency power unit mode, the engine 10 will bebrought up to speed more rapidly and, once at speed, will produce morepower. Typically, primary zone combustion temperatures of 3100° F.should be achieved within the combustor 48. After dilution of the gassesof combustion by the vaporized fuel, the temperature of the gasses tothe turbine will be approximately 1200° F. to 1800° F., well withinnormal operating limits for such a machine.

The system allows use of a fuel to air ratio of about 1 to 2 as opposedto about 1 to 14 or 15 for a stoichiometric reaction. Yet, approximatelythe same output in terms of horsepower per pound of fuel is obtainedmeaning that the system substantially reduces the amount of air requiredwhen the system is operating in the emergency power mode over that whichwould be required if so operating strictly on gasses of combustion.

In operation, the engine 10 is brought up to speed in the emergencypower unit mode or APU start mode in a very short period of time, lessthan four seconds. In many instances, even at high altitudes, once theengine 10 is at 90% of its operating speed, it may be switched out ofthe emergency power unit mode and into the auxiliary power unit mode toprovide electrical power since many turbine engines such as that shownat 10 can operate normally on the air available at high altitudes oncethey have been brought up to 90% or more of their operating speed.

It should also be observed that those components of the system thatenable it to function as an emergency power unit may be utilized instarting the system for use in an auxiliary power unit mode when on theground. In particular, and as seen in FIGS. 3 and 4, the air inlet 78 tothe combustor 48 may be connected via a valve 90 to an external sourceof air under pressure not shown. Using that source of air, the combustor48 may be fired and the engine 10 brought up to speed as mentionedpreviously. Once up to desired speed, the vanes 24 may be opened andcombustion initiated within the combustor 34, 36 to maintain the engine10 in operation in the auxiliary power unit mode.

This feature of the invention is highly desirable in a typical aircraftenvironment. In particular, while the aircraft is on the ground andprior to flight, there will typically be a requirement for use of thesystem in the auxiliary power unit mode. This, of course, requires thatthe engine 10 be started and utilizing the equipment required to causethe system to operate in an emergency power unit mode as a means ofstarting the engine 10, and then switching over to the auxiliary powerunit mode, operability of the system in both modes is necessarilychecked prior to flight. Further, note APU start reliability issignificantly enhanced with this system because even if ignition failureoccurred in the combustor 48, a normal cold air start will still beaccomplished.

Other features of the invention likewise contribute to its utility. Forexample, the installation of the second combustor 48 within the turbineexhaust housing conserves on space requirements as noted previously. Inaddition, because it is located in an area which already must beprovided with heat isolation insulation because of the hot temperaturesof the turbine exhaust, no other additional treatment for temperatureisolation need be made. And its installation or removal as a moduleprovides for ease of servicing.

Further, the use of the variable inlet guide vanes 24, which desirablywill be present in the turbine engine 10 in any event, to remove theload from the compressor 28 during start up and operation of theapparatus in the emergency power unit mode minimizes any need formechanical decoupling devices and the possibility of failure that goeswith them.

We claim:
 1. A method of generating emergency power with a power plantincludinga gas turbine engine including a centrifugal compressor coupledto a radial turbine; a generator coupled to said turbine engine to bedriven thereby; an air inlet to said compressor including at least onemovable vane for controlling air flow through said inlet; a firstreservoir for a combustible fuel; a first combustor for receiving andcombusting compressed air from said compressor and fuel from said firstreservoir to generate gasses of combustion and feeding said gasses ofcombustion to said turbine; a second reservoir for containing acompressed oxidant for combusting fuel from said first reservoir; asecond combustor for receiving and combusting fuel from said firstreservoir and oxidant from said second reservoir to generate gasses ofcombustion; means for introducing fuel from said first reservoir intothe gasses of combustion from said second combustor to vaporize suchfuel; and means for directing combined gasses of combustion andvaporized fuel to said turbine to drive the same; said method comprisingthe steps of(a) using said vanes (5) to substantially close said inletwhile (b) directing fuel and compressed oxidant to said secondcombustion to combust the same and (c) directing fuel to saidintroducing means.
 2. The method of claim 7 wherein steps (a), (b) and(c) are performed substantially simultaneously.
 3. The method of claim 8wherein the endurance of the performance of steps (a), (b) and (c) issufficiently long to bring said engine up to a speed whereat it mayoperate on gasses of combustion from said first combustor.
 4. The methodof claim 7 wherein step (b) is performed so as to combust fuel andcompressed oxidant approximately stoichiometrically.